CFD Results for an Axisymmetric Isentropic Relaxed Compression Inlet

The OVERFLOW code was used to calculate the flow field for a family of five relaxed compression inlets, which were part of a screening study to determine a configuration most suited to the application of microscale flow cont rol technology as a replacement for bleed. Comparisons are made to experimental data collected for each of the inlets in the 1'x1' supersonic wind tunnel at the NASA Glenn Research Center to help determine the suitability of CFD as a tool for future studi es of these inlets with flow control devices. Effects on the wind tunnel results of the struts present in a high subsonic flow region accounted for most of the inconsistency between the results. Based on the level of agreement in the present study, it is e xpected that CFD can be used as a tool to aid in the design of a study of this class of inlets with flow control. I. I ntroduction A supersonic inlet is used to decelerate and compress the flow before it enters the engine. Due to the compression, an advers e pressure gradient exists in the inlet that causes thick boundary layers to develop. Bleed has become the standard means of reducing the effects of the boundary layer. Part of the low velocity flow in the boundary layer is removed, leaving a higher averag e flow velocity in the inlet. However, since air is being removed from the system, the inlet must be larger to provide the same total mass flow to the engine, and the bleed flow is dumped overboard, which adds significant drag. Microscale flow control devi ces 1,2 are another potential means of controlling boundary layer development that are of interest as an alternative to bleed due to their low weight and mechanical simplicity. A series of five external compression, axisymmetric inlets were tested at the N ASA Glenn Research Center 1 ft x 1 ft supersonic wind tunnel 3 . Figure 1 shows a view of one of the inlet s in the tunnel. The inlets were part of a screening study to determine a configuration most suited to the application of microscale flow control techno logy as a replacement for bleed. Each inlet employed forward centerbody surface shaping to relax flow compression in the cowl highlight region. Relative to inlet geometry produced using traditional methods, this novel design technique can significantly red uce cowling angle and, therefore, nacelle drag. By capitalizing on fixed surface geometry redefinition, relaxed compression provides a simpler approach to improving installed propulsion performance compared to methods that focus on maximizing total pressur e recovery. While the novel shaping technique can provide a more mechanically simple approach to improving installed propulsion system performance, there are two primary limitations: tip distortion increases because of the stronger normal shock that result s from the relaxed isentropic compression, and boundary layer health is reduced behind the base of the terminal shock because of the increased post -shock flow turning angle required to maintain an acceptable subsonic diffusion profile. Based on the results of this first test, one of the inlets may be adapted to include microramps, which are wedge -shaped vortex generators with a height approximately 40% of the boundary layer